1. Field of the Invention
The present invention relates to the field of gas turbine engines such as turbojet engines.
2. Description of the Related Art
A multi-flow (for example dual-flow) turbojet engine, for the propulsion of commercial aircraft, generally comprises an upstream fan delivering an annular airflow, said flow comprising a primary central annular part, which supplies the motor driving the fan, in addition to a secondary external annular part, designed to be ejected into the atmosphere whilst providing a significant proportion of the thrust.
The airflow passing through the engine is compressed again to supply a combustion chamber, the gases thereof emerging from said chamber are then expanded in the turbine stages. To obtain the best possible efficiency, the gases leaving the chamber have to be at a high temperature, taking into account the behaviour of the materials forming the high pressure turbine immediately downstream of the combustion chamber. The high pressure turbine blades are cooled by the circulation of an airflow inside the aerofoil portion. Part of the cooling air is discharged through the wall of the bathtub-shaped cavity located at the distal end of said aerofoil portion.
Each blade comprises a pressure surface wall and a suction surface wall, in addition to a distal wall arranged at the distal end (at the apex) of the blade. Said walls are arranged so as to create, in the region of the distal end of the blade, at least one internal cavity and at least one external cavity, forming said bathtub-shaped cavity, separated from one another by the distal wall.
As the distal end of the pressure surface wall is particularly exposed, the apex of the blade is rapidly oxidized, which causes a loss of metal and thus eventually an increase in clearance in the region of the apex and, as a result, a reduction in the efficiency of the turbojet engine.
In order to cool the distal end of the pressure surface wall, it is already known to create therein openings for the flow of cooling air. More specifically, said openings take the form of holes made in the region of the pressure surface wall, not far from the distal wall and opening into the internal cavity.
Although such openings effectively permit the cooling of the extreme distal zone of the internal cavity of the blade, they prove to be located too far from the distal end of the pressure surface wall to cool it correctly.